Acoustic treatment in an unducted area of a geared turbomachine

ABSTRACT

An exemplary geared turbomachine assembly includes an acoustic treatment that is in an unducted area of a geared turbomachine.

This application is a continuation of U.S. application Ser. No.13/406,712, filed on Feb. 28, 2012.

BACKGROUND

This disclosure relates generally to acoustic treatments and, moreparticularly, to acoustic treatments in an unducted area of a gearedturbomachine.

Turbomachines, such as gas turbine engines, typically include a fansection, a turbine section, a compressor section, and a combustorsection. Turbomachines may employ a geared architecture that allows thefan to spin at a slower rotational speed than the low pressure turbine.

Air moves into the turbomachine through the fan section. Some of thisair moves into a core of the turbomachine. The remaining air movesthrough a bypass flowpath established between a fan cowl and a coreengine cowl of the core. The core of the turbomachine extends axiallyoutside the fan cowl. That is, some of the bypass stream is ducted andsome of it is unducted.

Acoustic treatments attenuate noise radiating from the turbomachine.These acoustic treatments are traditionally limited to ducted areas ofthe turbomachine. Extending acoustic treatments to unducted areas isless efficient as a result of the reduced interaction between thepropagating noise and the acoustic treatment surface. Moreover, placingthe acoustic treatment in unducted areas subject to high Mach numberflows (e.g. near a nozzle exit) results in a performance penaltyassociated with increased drag. In a typical turbofan with a fanpressure ratio greater than 1.5, the high performance penalty associatedwith the high nozzle Mach number along with the reduced noise benefitmakes it impractical to place acoustic treatment on the unducted areapast the fan nozzle exit plane.

SUMMARY

A geared turbomachine assembly according to an exemplary aspect of thepresent disclosure includes, among other things, an acoustic treatmentthat is in an unducted area of a geared turbomachine.

In a further non-limiting embodiment of the foregoing gearedturbomachine assembly, a portion of the acoustic treatment may beaxially aft a fan cowl of the geared turbomachine.

In a further non-limiting embodiment of either of the foregoing gearedturbomachine assemblies, a portion of the acoustic treatment is axiallyaligned with a low-pressure turbine section of the geared turbomachine.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may extend continuouslyfrom the ducted area to the portion that is axially aligned with thelow-pressure turbine section of the geared turbomachine.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may form a portion of acore engine cowl.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may extend axially to anaftmost end of the core engine cowl.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may extend continuouslyfrom the ducted area to the aftmost end.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may comprise aperforated face sheet having an open area density that is from 4% to30%.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may comprise amicroperforated face sheet having apertures with a diameter that is lessthan 25 percent of the perforated face sheet.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the geared turbomachine assembly may have a fanpressure ratio that is less than 1.45 to 1.

A geared turbomachine assembly according to another exemplary aspect ofthe present disclosure includes, among other things, a fan cowl housinga fan that is driven by a geared architecture, a core cowl housing acore engine, an acoustic treatment establishing a portion of the corecowl aft the fan cowl.

In a further non-limiting embodiment of the foregoing gearedturbomachine assembly, the geared turbomachine assembly may have aportion of the acoustic treatment axially aft a fan cowl of the gearedturbomachine.

In a further non-limiting embodiment of either of the foregoing gearedturbomachine assemblies, a portion of the acoustic treatment may beaxially aligned with a low-pressure turbine section of the gearedturbomachine.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may form a portion ofthe core cowl.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may extend axially to anaftmost end of the core cowl.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may comprise aperforated face sheet having an open area density that is from 4% to30%.

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the acoustic treatment may comprise amicroperforated face sheet having apertures with a diameter that is lessthan 25 percent of the perforated face sheet

In a further non-limiting embodiment of any of the foregoing gearedturbomachine assemblies, the geared turbomachine assembly may have a fanpressure ratio that is less than 1.45 to 1.

A method of attenuating noise in a geared turbomachine according toanother exemplary aspect of the present disclosure includes, among otherthings, attenuating noise axially aft of a fan cowl using an acoustictreatment that provides a portion of an engine core, wherein the fancowl houses a fan driven by a geared architecture.

In a further non-limiting embodiment of the foregoing method ofattenuating noise in a geared turbomachine, the acoustic treatment mayextend axially past a high-pressure turbine section.

DESCRIPTION OF THE FIGURES

The various features and advantages of the disclosed examples willbecome apparent to those skilled in the art from the detaileddescription. The figures that accompany the detailed description can bebriefly described as follows:

FIG. 1 shows a schematic view of an example turbomachine having a gearedarchitecture.

FIG. 2 shows a cross-section view of a portion of another exampleturbomachine having a geared architecture.

FIG. 3 shows a partial section view of an example acoustic treatmentused in the FIG. 2 engine.

FIG. 4 shows a partial section view of another example acoustictreatment used in the FIG. 2 engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example turbomachine, which is a gasturbine engine 20 in this example. The gas turbine engine 20 is atwo-spool turbofan gas turbine engine that generally includes a fansection 22, a compressor section 24, a combustion section 26, and aturbine section 28. Other examples may include an augmentor section (notshown) among other systems or features.

Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with turbofans. Thatis, the teachings may be applied to other types of turbomachines andturbine engines including three-spool architectures.

The example engine 20 generally includes a low-speed spool 30 and ahigh-speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36. Thelow-speed spool 30 and the high-speed spool 32 are rotatably supportedby several bearing systems 38. It should be understood that variousbearing systems 38 at various locations may alternatively, oradditionally, be provided.

The low-speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low-pressure compressor 44, and a low-pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than thelow-speed spool 30.

The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh-pressure compressor 52 and high-pressure turbine 54.

The inner shaft 40 and the outer shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A, whichis collinear with the longitudinal axes of the inner shaft 40 and theouter shaft 50.

The combustion section 26 includes a circumferentially distributed arrayof combustors 56 generally arranged axially between the high-pressurecompressor 52 and the high-pressure turbine 54.

In some non-limiting examples, the engine 20 is a high-bypass gearedaircraft engine. In a further example, the engine 20 bypass ratio isgreater than about six (6 to 1).

The geared architecture 48 of the example engine 20 includes anepicyclic gear train, such as a planetary gear system or other gearsystem. The example epicyclic gear train has a gear reduction ratio ofgreater than about 2.3 (2.3 to 1).

The low-pressure turbine 46 pressure ratio is pressure measured prior toinlet of low-pressure turbine 46 as related to the pressure at theoutlet of the low-pressure turbine 46 prior to an exhaust nozzle of theengine 20. In one non-limiting embodiment, the bypass ratio of theengine 20 is greater than about ten (10 to 1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low-pressure turbine 46 has a pressure ratio that is greater thanabout 5 (5 to 1). The geared architecture 48 of this embodiment is anepicyclic gear train with a gear reduction ratio of greater than about2.5 (2.5 to 1). It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a geared architectureengine and that the present disclosure is applicable to other gasturbine engines including direct drive turbofans.

In this embodiment of the example engine 20, a significant amount ofthrust is provided by the bypass flow B due to the high bypass ratio.

The fan section 22 of the engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet. Thisflight condition, with the engine 20 at its best fuel consumption, isalso known as “Bucket Cruise” Thrust Specific Fuel Consumption (TSFC).TSFC is an industry standard parameter of fuel consumption per unit ofthrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The FanPressure Ratio according to one non-limiting embodiment of the exampleengine 20 is less than 1.45 (1.45 to 1), which is relatively low.

Low Corrected Fan Tip Speed is the actual fan tip speed divided by anindustry standard temperature correction of Temperature divided by 518.7A 0.5. The Temperature represents the ambient temperature in degreesRankine. The Low Corrected Fan Tip Speed according to one non-limitingembodiment of the example engine 20 is less than about 1150 fps (351m/s).

The example gas turbine engine 20 generally includes a ducted area 60and an unducted area 62. In this example, the ducted area 60 representsthe portions of the gas turbine engine 20 axially aligned with, andradially bounded by, a fan cowl 66 (or nacelle). The unducted arearepresents the portions of the gas turbine engine 20 axially outside ofthe fan cowl 66.

In this example, a core engine cowl 70 of the gas turbine engine 20includes portions in the ducted area 60 and portions in the unductedarea 62. The core engine cowl 70 includes an acoustic treatment 74 thatextends continuously from the ducted area 60 to the unducted area 62.The acoustic treatment 74 is on an outer surface of the core engine cowl70. That is, at least some of the acoustic treatment 74 is in anunducted area of the gas turbine engine 20.

Referring to FIG. 2, another example geared turbomachine is the gasturbine engine 120. In this example, the acoustic treatment 74 extendsfrom the ducted area 160, past an aft end 76 of the fan cowl 66, to theunducted area 162 of the engine 120. Notable, the example acoustictreatment 74 extends continuously, and without interruption, from theducted area 160 to the aftmost end 84 of a cowl 80 of the core enginecowl 170.

A portion of the acoustic treatment 74 is axially aligned with theturbine section 128 of the engine 120, and, in particular, thelow-pressure turbine 146 within the turbine section 128. The acoustictreatment 74 is continuous and uninterrupted from the ducted area 160,past the high-pressure turbine 154, to the portion that is aligned withthe low-pressure turbine 146. In this example, the acoustic treatment 74also lines radially inwardly facing surfaces of the fan cowl 166.

Referring to FIG. 3 with continuing reference to FIG. 2, an exampleacoustic treatment 74 includes a perforated face sheet 78 and a backingsheet 82. A honeycomb structure 86 is sandwiched between the perforatedface sheet 78 and the backing sheet 82. The perforated face sheet 78,the backing sheet 82 and the honeycomb structure are all aluminum inthis example. Other examples of the acoustic treatment 74 may be made ofa composite material or some other material.

The example perforated face sheet 78 includes an array of perforations88 each having a diameter that is from 0.03 inches (0.762 mm) to 0.06inches (1.524 mm). A density of the open area in the perforated facesheet 78 is from 6% to 20% in this example. Other perforated face sheets78 have perforation 88 having other diameters.

In this example, the core engine cowl 170 in the areas of the acoustictreatment 74 is formed entirely by the acoustic treatment 74. That is,there is no walled structure that the acoustic treatment 74 is securedto. The acoustic treatment 74 is at an outer surface of the core enginecowl 170.

In another example, the acoustic treatment 74 and, in particular, theback sheet of the acoustic treatment 74, is adhesively secured to a wall(not shown). In such an example, the adhesively secured acoustictreatment and the wall together form the core engine cowl 170.

The acoustic treatment 74 is a perforated acoustic treatment with asingle degree of freedom (SDOF) in this example, other examples mayinclude a Double Degree Of Freedom (DDOF) liner, bulk absorber liner ora microperforated acoustic treatment 174 as shown in FIG. 4. Themicroperforated acoustic treatment 174 includes a perforated face sheet178 and a backing sheet 182 sandwiching a honeycomb structure 186.Perforations 188 in the perforated face sheet 178 each have a diameterthat is less than 25 percent of a thickness t of the perforated facesheet 178.

Features of the disclosed examples include incorporating an acoustictreatment in an area of a turbomachine that is traditionally not wellsuited for acoustic treatments due to the associated drag penalties. Theadded acoustic treatment has minimal impact on fuel burn.

In this disclosure, like reference numerals designate like elementswhere appropriate, and reference numerals with the addition ofone-hundred or multiples thereof designate modified elements. Themodified elements incorporate the same features and benefits of thecorresponding modified elements, expect where stated otherwise.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. Thus, the scope of legal protectiongiven to this disclosure can only be determined by studying thefollowing claims.

1. A geared turbomachine assembly, comprising: an acoustic treatmentestablishing a portion of a core cowl aft a fan cowl, the acoustictreatment including a perforated face sheet, a backing sheet, and acellular structure extending therebetween, the perforated face sheethaving an open area density that is from 4 percent to 30 percent. 2.(canceled)
 3. The geared turbomachine assembly of claim 1, wherein aportion of the acoustic treatment is axially aligned with a low-pressureturbine section of the geared turbomachine.
 4. The geared turbomachineassembly of claim 3, wherein the acoustic treatment extends from aposition that is axially aligned with the fan cowl to the portion thatis axially aligned with the low-pressure turbine section of the gearedturbomachine.
 5. (canceled)
 6. The geared turbomachine assembly of claim1, wherein the acoustic treatment extends axially to an aftmost end ofthe core engine cowl.
 7. The geared turbomachine assembly of claim 6,wherein the acoustic treatment extends continuously from a position thatis axially aligned with the fan cowl to the aftmost end.
 8. (canceled)9. The geared turbomachine assembly of claim 1, wherein the perforatedface sheet comprises a microperforated face sheet having apertures witha diameter that is less than 25 percent of a thickness of themicroperforated face sheet.
 10. The geared turbomachine assembly ofclaim 1, wherein the geared turbomachine assembly has a fan pressureratio less than 1.45.
 11. A geared turbomachine assembly, comprising: afan cowl housing a fan that is driven by a geared architecture; a corecowl housing a core engine; and an acoustic treatment establishing aportion of the core cowl aft the fan cowl, the acoustic treatmentcomprising a cellular structure sandwiched between a microperforatedface sheet and a backing sheet, the microperforated face sheet includingapertures having a diameter that is less than 25 percent of a thicknessof the microperforated face sheet.
 12. The geared turbomachine assemblyof claim 11, wherein a portion of the acoustic treatment is axially aftthe fan cowl of the geared turbomachine assembly.
 13. The gearedturbomachine assembly of claim 11, wherein a portion of the acoustictreatment is axially aligned with a low-pressure turbine section of thegeared turbomachine assembly.
 14. The geared turbomachine assembly ofclaim 11, wherein the acoustic treatment forms a portion of the corecowl.
 15. The geared turbomachine assembly of claim 14, wherein theacoustic treatment extends axially to an aftmost end of the core cowl.16. The geared turbomachine assembly of claim 11, wherein themicroperforated face sheet has an open area density that is from 4% to30%.
 17. (canceled)
 18. The geared turbomachine assembly of claim 11,wherein the geared turbomachine assembly has a fan pressure ratio lessthan 1.45.
 19. A method of attenuating noise in a geared turbomachine,comprising: attenuating noise axially aft of a fan cowl using anacoustic treatment that provides a portion of an engine core cowl,wherein the fan cowl houses a fan driven by a geared architecture, theacoustic treatment comprising a cellular structure sandwiched between amicroperforated face sheet and a backing sheet, the microperforated facesheet including apertures having a diameter that is less than 25 percentof a thickness of the microperforated face sheet.
 20. The method ofclaim 19, wherein the acoustic treatment extends axially pastahigh-pressure turbine section.
 21. A method of configuring aturbomachine to attenuate noise, the turbomachine including a fan cowlhousing a fan driven by a geared architecture; the method comprising:providing an acoustic treatment to a portion of an engine core cowlaxially aft of the fan cowl, the acoustic treatment including aperforated face sheet, a backing sheet, and a cellular structureextending therebetween, the perforated face sheet having an open areadensity that is from 4 percent to 30 percent.
 22. The method of claim21, wherein the acoustic treatment extends axially pasta high-pressureturbine section.
 23. The geared turbomachine assembly of claim 1,wherein the core cowl is formed entirely by the acoustic treatment insome axial areas.
 24. The geared turbomachine assembly of claim 1,wherein a selected axial portion of the acoustic treatment is notsecured to a walled structure.
 25. The geared turbomachine assembly ofclaim 11, wherein a selected axial portion of the core cowl is formed bythe acoustic treatment.
 26. The geared turbomachine assembly of claim11, wherein the microperforated face sheet has an open area density thatis from 4% to 30%.